Inertial reference system for an aircraft

ABSTRACT

An inertial reference system for an aircraft includes two accelerometers and a gyrometer. One accelerometer is located at a front portion of the aircraft, the other is located at a rear portion of the aircraft. The gyrometer is located at a center portion of the aircraft. The system also includes a control computer linked to the accelerometers and to the gyrometer. The center portion can include the aircraft&#39;s center of gravity.

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] This application is a continuation application of published U.S.application Ser. No. 10/336,730, filed Jan. 6, 2003, which claimspriority under 35 U.S.C. § 119 to French Patent Application 02 04334,filed on Apr. 8, 2002, the entire disclosure of both which areincorporated herein by reference.

BACKGROUND OF THE INVENTION

[0002] 1. Field of the Invention

[0003] The present invention relates to aircraft with electric flightcontrols including a fuselage able to deform and vibrate longitudinallyand laterally with the formation of vibration nodes and antinodesdistributed along the longitudinal axis of the aircraft. It relatesquite particularly to long-length airplanes which have high longitudinalflexibility. However, it advantageously applies equally well toairplanes of a shorter length and lower flexibility.

[0004] 2. Discussion of the Background

[0005] It is known that an aircraft with electric flight controls hasflight controls such as sticks, mini sticks, rudder bars, etc., whichare equipped with electric transducers so that they generate electricflight control datums representative of the action that a pilot exertson them. It also includes a flight control computer which, on the basisof the electric flight control datums generated by the flight controlsand of flight control parameters originating, for example, from sensors,formulates electric commands that the flight control computer applies toactuators tasked with moving the control surfaces of the aircraft.

[0006] It is also known that aircraft with electric flight controls areprovided with an inertial reference system (generally known as an IRS)including elements useful in navigation, such as the inertial unit, andelements useful in flight control, such as gyrometers andaccelerometers. Finally, it is known that all these elements, whetherthey have to do with navigation or flight control, are grouped togetherin an IRS unit arranged at a given point on the aircraft. Of course, asa result, this IRS unit is subjected to the action of the deformationsof the fuselage, which deformations occur mainly along the axes of pitchand yaw under the effect of the turning of the control surfaces or theeffect of external disturbances.

[0007] Because of the high time constant attached to the elements usefulin navigation, such deformations have only a small action thereon. Bycontrast, in order to get around the problems of interaction between thedeformations of the fuselage and the elements useful in flight control,it is essential to have filtering means on the control surface controllines.

[0008] However, in the case of aircrafts with high longitudinalflexibility, the deformations become greater, which means that it isthen necessary to perform extremely intense filtering of the controllines, and this introduces significant phase shifts thereinto andtherefore detracts greatly from the performance of the control lines.

SUMMARY OF THE INVENTION

[0009] It is an object of the present invention to overcome thisdrawback.

[0010] To this end, according to the invention, an aircraft withelectric flight controls, provided with control surfaces able to bemoved by electrically operated actuators, includes controls and at leastone flight control computer. The controls are actuated by a pilot andgenerate electric flight control datums which are sent to the flightcontrol computer. The latter computer generates, on the basis of theelectric flight control datums and flight control parameters, commandsin roll, pitch and yaw, which are sent to the actuators to move thecontrol surfaces. An inertial reference system includes elements usefulin navigation and elements useful in flight control, the latter elementsbeing either of the gyrometer type or the accelerometer type. Theaircraft includes a fuselage able to deform and vibrate with theformation of vibration nodes and antinodes distributed along thelongitudinal axis of the aircraft.

[0011] The inertial reference system has an exploded structure with theelements useful in flight control separated from the elements useful innavigation. The elements useful in flight control are distributed alongthe fuselage. Each element useful in flight control, of the gyrometertype, is arranged at a vibration node of the fuselage. Each elementuseful in flight control, of the accelerometer type, is arranged at avibration antinode of the fuselage. The elements useful in flightcontrol are connected to the flight control computer so that themeasurement signals they deliver are used as flight control parameters.

[0012] Thus, the accelerometers allow the measurement of theaccelerations of the aircraft including vibrational movements of thefuselage, while the gyrometers allow the measurement of the rotationrates without incorporating the structural modes of the fuselagethereinto. These accelerometer and gyrometer measurements are sent tothe flight control computer which in consequence formulates commands forthe control surfaces.

[0013] The flight control laws incorporated into this computer thereforedo not need to filter the vibrational movements of the fuselage. This isbecause the structural modes measured by the accelerometers can beactively checked by the flight control laws while the gyrometers do notmeasure deformations of the fuselage. In the most frequent scenario, theaircraft fuselage deforms and vibrates in such a way as to have avibration antinode at each of its ends, and a vibration node near itscenter of gravity.

[0014] In this case, the aircraft includes at least one frontaccelerometer arranged at the front part of the fuselage and deliveringa vertical acceleration measurement and a lateral accelerationmeasurement. At least one rear accelerometer is arranged at the rearpart of the fuselage and delivering a vertical acceleration measurementand a lateral acceleration measurement. At least one gyrometer isarranged near the center of gravity of the aircraft and delivering rollrate, pitch rate and yaw rate measurements.

[0015] It is then advantageous for the vertical accelerationmeasurements generated by the front accelerometer and by the rearaccelerometer respectively and the pitch rate measurement generated bythe gyrometer, to be used as flight control parameters to formulate thepitch commands. The lateral acceleration measurements generated by thefront accelerometer and by the rear accelerometer respectively, and theroll rate and yaw rate measurements generated by the gyrometer, can beused as flight control parameters to formulate the roll commands. Thelateral acceleration measurements generated by the front accelerometerand by the rear accelerometer respectively, and the roll rate and yawrate measurements generated by the gyrometer, can be used as flightcontrol parameters for formulating the yaw commands.

[0016] The aircraft can include means of filtering the accelerationmeasurements and the rate measurement or measurements to eliminatemeasurement noise therefrom and avoid spectrum folding. The aircraft canalso include gain multipliers for weighting each of the filteredacceleration or rate measurements; phase control means for the filteredand weighted acceleration measurements; and summing means for summingthe filtered, weighted and phase-controlled acceleration measurements,the filtered and weighted rate measurement or measurements and thecorresponding electric flight control datum to formulate thecorresponding command.

[0017] The aircraft may also, for formulating roll and yaw commands,include means of integrating the roll rate so as to create informationabout the roll angle, which information is sent to the summing meansafter it has been weighted by a gain multiplier.

[0018] Of course, in such an architecture, all the gains are optimizedso as to satisfy the compromises between performance and stability. Itis also found that the architecture according to the present inventionmakes it possible to dispense with low-frequency filters, even thoughthe aircraft might be very flexible.

BRIEF DESCRIPTION OF THE DRAWINGS

[0019] The figures of the appended drawing will make it easier tounderstand how the invention may be embodied. In these figures,identical references denote similar elements.

[0020]FIG. 1 schematically and generally illustrates the electric flightcontrol system according to the present invention, the one example of anairplane with high longitudinal flexibility.

[0021]FIG. 2 shows, in schematic perspective, a civil transportairplane, with the locations of its accelerometers and gyrometers.

[0022]FIG. 3 is the block diagram of the pitch control system of theairplane of FIG. 2.

[0023]FIG. 4 is the block diagram of the roll and yaw control systems ofthe airplane of FIG. 2.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0024] The airplane 1 with high flexibility along its longitudinal axisL-L, shown in FIG. 1, can deform under the effect of the turning of itscontrol surfaces or of external disturbances so that the maindeformation of its fuselage 2 in the yaw and pitch axes is verysignificant at the front 3AV and rear 3AR ends of the fuselage 2 whilethe center 4 of this fuselage (at which the center of gravity of theairplane 1 is located) deforms little. In addition, the rotation ratesassociated with the deformations of the fuselage 2 are very small nearthe center 4 of the fuselage.

[0025] As illustrated schematically in FIG. 1, the airplane 1 includes:

[0026] an inertial unit CI, intended for navigation and arranged at anycustomary and appropriate point on the fuselage 2;

[0027] at least one front accelerometer 5 arranged at the front end 3AV;

[0028] at least one rear accelerometer 6 arranged at the rear end 3AR;and

[0029] at least one gyrometer 7 near the center 4 of the airplane 1.

[0030] Of course, although in FIG. 1 the accelerometers 5 and 6 and thegyrometer 7 are depicted on the outside of the airplane 1 to make thedrawings clear, they are, in actual fact, housed inside the fuselage 2as depicted schematically in FIG. 2.

[0031] The front and rear accelerometers 5 and 6 make it possible tomeasure the accelerations of the airplane 1, including the vibrationalmovements of the fuselage 2, these accelerations being measured in theform of their lateral components (NYAV in the case of the frontaccelerometer 5, and NYAR in the case of the rear accelerometer 6) andvertical components (NZAV in the case of the front accelerometer 5, andNZAR in the case of the rear accelerometer 6). Moreover, the gyrometer 7makes it possible to measure the rotation rates of the fuselage 2 nearthe center of gravity of the airplane 1, excluding the contribution ofthe structural modes thereof. These rotation rates are broken down intotheir three components P (roll rate), Q (pitch rate) and R (yaw rate)near the center of gravity of the airplane 1.

[0032] Moreover, the airplane 1 includes:

[0033] at least one stick 8, for example of the mini stick type, intendsto be actuated by a pilot (not depicted) and associated with atransducer 9 generating roll and pitch flight control datumsrepresentative of the movements of the stick 8;

[0034] at least one rudder bar 10 intended to be actuated by the pilotand associated with a transducer 11 generating yaw flight control datumsrepresentative of the movements of the rudder bar 10;

[0035] at least one flight control computer 12 which, in the usual way,receives:

[0036] via links 13, the roll and pitch flight control datums generatedby the controls 8, 9;

[0037] via links 14, the yaw flight control datums generated by thecontrols 10, 11; and

[0038] via links 15, flight control parameters originating from sensors,other computers, etc.

[0039] Some of the links 15 connect the accelerometers 5 and 6 and thegyrometer 7 to the flight control computer 12 so that the measurementsNZAV, NZAR, NYAV, NYAR, P, Q and R form part of the flight controlparameters sent to the computer 12 via the links 15.

[0040] On the basis of the roll, pitch and yaw flight control datums andof the flight control parameters, the flight control computer 12generates commands which are sent to a number of actuators 16.1, 16.2, .. . , 16.i, . . . , 16.n each of which moves a control surface 17.1,17.2, . . . , 17.i, . . . , 17.n accordingly.

[0041] It can be seen that the structural vibration modes measured bythe accelerometers 5 and 6 can thus be actively checked by the flightcontrol laws embedded in the computer 12, while the gyrometer 7 does nottake fuselage deformation into consideration. There is therefore noneed, using these flight control laws, to filter the vibrationalmovements of the fuselage 2.

[0042] As can be seen in FIG. 2, the accelerometers 5 and 6 are arrangedrespectively at locations 18 and 19 at the front end 3AV and at the rearend 3AR of the airplane 1. Furthermore, the airplane includes:

[0043] an elevator 21, articulated to the trailing edge of an adjustablehorizontal plane 22;

[0044] ailerons 23 and spoilers 24, articulated to the trailing edge ofthe wings 25; and

[0045] a rudder 26 articulated to the trailing edge of the verticalstabilizer 27.

[0046] Of course, each of these control surfaces 21 to 24 and 26corresponds to one of the control surfaces 17.i (where i=1 to n) in FIG.1.

[0047]FIG. 3 schematically depicts the part 12A of the flight controlcomputer 12 corresponding to pitch control in accordance with thepresent invention and intended to control the elevator 21 and theadjustable horizontal plane 22. This control is effected through frontand rear vertical acceleration measurements NZAV and NZAR and themeasurement of the pitch rate Q near the center 4, which are sent to itvia the corresponding links 15.

[0048] In this part 12A of the flight control computer 12, eachmeasurement NZAV, NZAR and Q is filtered by respective filter means 28,29 and 30, and weighted with a gain, by gain multipliers 31, 32 and 33respectively. Such filtering, the purpose of which is to avoid noise andspectrum folding, relates to the high frequencies in excess of 10 Hz. Itis therefore not penalizing to the performance of the pitch control. Inaddition, phase controllers 34 and 35 receiving the weightedaccelerometer measurements NZAV and NZAR are able actively to check thestructural modes of the fuselage 2. Such phase control corresponds to anadjustment of the pitch control law, the adjustment being pegged to thephase of the structural modes, so as to increase their damping. Thesignals leaving the phase controllers 34 and 35 and the gain multiplier33 are summed in a summer 36, making it possible at output therefrom toobtain a pitch command that is a function of the three measurementsNZAV, NZAR and Q.

[0049] Furthermore, this part 12A of the computer 12 additionallyincludes a processing device 37 and a gain multiplier 38 for the pitchflight control datum generated by a control 8, 9 and sent to the device37 via a link 13.

[0050] This pitch flight control datum thus processed and weighted bythe device 37 and the multiplier 38 sent to a summer 39 in which it issummed with the pitch command that appears at output from the summer 36.

[0051] The composite pitch command appearing at the output of the summer39 is sent to the actuators 16.i of the elevator 21 and of theadjustable horizontal plane 22 to move these accordingly.

[0052]FIG. 4 schematically depicts the parts 12B and 12C of the flightcontrol computer 12 correspondingly respectively to roll control bymeans of the ailerons 23 and the spoilers 24 and to yaw control by meansof the rudder 26. These two parts 12B and 12C of the computer 12receive, via the corresponding links 15, the lateral accelerationmeasurements NYAV and NYAR delivered by the accelerometers 5 and 6,together with the roll rate P and yaw rate R which are measured by thegyrometer 7.

[0053] In each of the parts of the computer 12B and 12C, eachmeasurement NYAV, NYAR, P and R is filtered by high-frequency filteringmeans (frequency in excess of 10 Hz) 40, 41; 42, 43; 44, 45; 46, 47,respectively, allowing the corresponding commands to get around problemsof noise and spectrum folding without disadvantageous influence on theperformance of the commands. In addition, the measurements are weightedusing gains, by virtue of respective gain multipliers 48, 49; 50, 51;52, 53; 54, 55. Respective phase controllers 56, 57 and 58, 59(analogous to the controllers 34 and 35 of the part 12A of the computer12) receive the weighted accelerometer measurements NYAV and NYAR so asto check actively the structural modes of the fuselage 2. The signalsleaving the controllers 56 and 58 and the gain multipliers 52 and 54 aresent to summers 60. Likewise, the signals leaving the controllers 57 and59 and the gain multipliers 53 and 55 are sent to a summer 61.

[0054] In addition, in each part of the computer 12B or 12C, thefiltered roll rate P appearing at the outputs of the filtering means 44or 45 respectively is integrated by an integrator 62 or 63 then weightedby a gain multiplier 64 or 65. Such integration actions make it possibleto create information about the roll angle, which information is sent tothe respective summer 60 or 61.

[0055] Thus, at the outputs from the summers 60 and 61 there areobtained, respectively, a roll command and a yaw command each of whichis a function of the four measurements NYAV, NYAR, P and R and of theroll angle information resulting from integration by the integrator 62or 63 respectively.

[0056] The flight computer part 12B additionally includes a processingpart 62 and a gain multiplier 64 for the roll flight control datumgenerated by a flight control 8, 9 and sent to the device 62 by a link13. This roll flight control datum thus processed and weighted by thedevice 62 and the gain multiplier 64 is sent to a summer 66 in which itis summed with the roll command appearing at the output of the summer60. The composite roll command appearing at the output of the summer 66is sent to the actuators 16.i of the ailerons 23 and of the spoilers 24.

[0057] Likewise, the part of the computer 12C additionally includes aprocessing device 63 and a gain multiplier 65 for the yaw flight controldatum generated by a flight control 10, 11 and sent to the device 63 bya link 14. This yaw flight control datum thus processed and weighted bythe device 63 and the gain multiplier 65 is sent to a summer 67 in whichit is summed with the yaw command appearing at the output of the summer61. The composite yaw command appearing at the output of the summer 67is sent to the actuators 16.i of the rudder 26.

What is claimed is:
 1. An inertial reference system for an aircraft, comprising: a first accelerometer located at a front portion of said aircraft; a second accelerometer located at a rear portion of said aircraft; a gyrometer located at a center portion of said aircraft; and a control computer linked to said first and second accelerometers and to said gyrometer.
 2. The system of claim 1, wherein the center portion include said aircraft's center of gravity.
 3. The system of claim 1, wherein said control computer is located outside said front, rear, and central portions.
 4. The system of claim 1, wherein said first and second accelerometers and said gyrometer are housed inside a fuselage of said aircraft.
 5. The system of claim 1, wherein said control computer generates flight control parameters based on data received from said first and second accelerometers and to said gyrometer.
 6. A system for controlling an aircraft, comprising means for receiving first vertical acceleration data related to a vertical acceleration of a front portion of said aircraft; means for receiving second vertical acceleration data related to a vertical acceleration of a rear portion of said aircraft; means for receiving pitch rate data related to a pitch rate of a center portion of said aircraft; and means for generating a pitch command based on said first and second vertical acceleration data and on said pitch rate data.
 7. The system of claim 6, wherein said means for generating comprises means for filtering signals carrying said first and second vertical acceleration data and said pitch rate data.
 8. The system of claim 7, wherein said means for filtering filters frequencies in excess of 10 Hz.
 9. The system of claim 6, further comprising means for receiving pitch flight control data, and wherein said means for generating generates said pitch command based on said pitch flight control data.
 10. A system for controlling an aircraft, comprising means for receiving first horizontal acceleration data related to a horizontal acceleration of a front portion of said aircraft; means for receiving second horizontal acceleration data related to a horizontal acceleration of a rear portion of said aircraft; means for receiving roll rate data related to a roll rate of a center portion of said aircraft; means for receiving yaw rate data related to a yaw rate of a center portion of said aircraft; and means for generating at least one of a roll command and a yaw command based on said first and second horizontal acceleration data, on said roll rate data, and on said yaw rate data.
 11. The system of claim 10, wherein said means for generating comprises means for filtering signals carrying said first and second horizontal acceleration data, said roll rate data and said yaw rate data.
 12. The system of claim 10, wherein said means for filtering filters frequencies in excess of 10 Hz.
 13. The system of claim 10, further comprising means for receiving roll flight control data, wherein said means for generating generates said roll command based on said roll flight control data.
 14. The system of claim 10, further comprising means for receiving yaw flight control data, wherein said means for generating generates said yaw command based on said yaw flight control data.
 15. The system of claim 10, further comprising: means for receiving first vertical acceleration data related to a vertical acceleration of said front portion of said aircraft; means for receiving second vertical acceleration data related to a vertical acceleration of said rear portion of said aircraft; means for receiving pitch rate data related to a pitch rate of said center portion of said aircraft; and means for generating a pitch command based on the first and second vertical acceleration data and on the pitch rate data. 